(1) Field of the Invention
The present invention relates to an aircraft servo-control provided with a limit force detector device, said aircraft being constituted by a rotorcraft and in particular by a helicopter.
(2) Description of Related Art
Conventionally, an aircraft includes control members that are capable of being controlled by the pilot, such as the blades of a helicopter lift rotor, or even airplane rudders, for example.
Using flight controls, the pilot causes control members of the aircraft to move. Nevertheless, the forces that need to be delivered in order to move the control members are sometimes very great.
Consequently, the linkage connecting a flight control to a control member is often fitted with a servo-control enabling the pilot to control the aircraft accurately and without difficulty.
More particularly, a helicopter is provided with a main rotor providing it with lift and propulsion. In order to control the helicopter, a pilot changes the pitch of the blades of the main rotor, i.e. their angle of incidence relative to the incident stream of air.
As a result, the rotorcraft includes a cyclic swashplate having a stationary bottom plate and a rotary upper plate. The stationary lower plate is connected to the pilot's flight controls, generally via three distinct control lines, while the rotary upper plate is connected to each of the blades via a respective rod. The cyclic swashplate is thus a control device that slides vertically along the mast of the main rotor to control the general pitch of the blades of the main rotor while also being capable of oscillating in any direction about a ball joint so as to control the cyclic pitch of the blades.
The oscillations and the vertical movements of the cyclic swashplate as controlled by the pilot thus cause the pitch of the blades to vary and they enable the pilot to control the helicopter.
Conventionally, the pilot controls the cyclic swashplate via mechanical controls that are connected to said cyclic swashplate by rods. Nevertheless, the forces that the pilot needs to exert in order to move the cyclic swashplate are very large, in particular if the weight of the rotorcraft is relatively great.
Consequently, a servo-control is arranged between an upstream portion and a downstream portion of each control linkage. The pilot then acts on the servo-controls via the upstream portion of the linkage and without exerting large amounts of force, and then the servo-controls reproduce the pilot's orders by delivering greater amounts of force to the downstream portion of the linkage.
Similarly, a helicopter is provided with a tail rotor and the pitch of its blades can be modified via a servo-control.
Naturally, the same applies for example to the ailerons of airplanes that are operated via servo-controls.
It should be observed that some modern aircraft include electric flight controls that replace the mechanical connections that are used to connect the flight controls to the servo-controls.
In conventional manner, the servo-controls include at least an outer body of cylindrical shape having a slider element movable in translation therein and provided with a rod carrying a control piston, the control piston defining a retraction chamber and an extension chamber inside the outer body.
The servo-control may also include a hydraulic directional control valve or it may co-operate with such a hydraulic control valve. Movement of the control piston on the slider element relative to the outer body is then controlled by the hydraulic control valve, which is itself actuated by the flight controls of the helicopter pilot via the upstream portion of a linkage. Depending on the orders received, the hydraulic control valve feeds hydraulic fluid to the retraction chamber or to the extension chamber in order to cause the servo-control to retract or to extend.
It will be understood that in the text below the term “retraction chamber” is used to designate a chamber causing the servo-control to retract when said chamber is filled with a fluid. Conversely, the term “extension chamber” is used to designate a chamber that causes the servo-control to extend when said chamber is filled with a fluid.
Two embodiments of servo-controls then coexist.
In a first embodiment, the control piston is secured to a stationary point of the aircraft, with the body moving as a function of the orders received and being connected to the downstream portion of the linkage. The person skilled in the art may refer to this type of servo-control as a “moving body servo-control”.
In contrast, in a second embodiment, the body is secured to a stationary point of the aircraft, with the control piston moving as a function of the orders received and being connected to the downstream portion of the linkage. The person skilled in the art may refer to this type of servo-control as a “stationary body servo-control”.
Furthermore, whatever the embodiment, there are servo-controls that the person skilled in the art may refer to as “single-body” or as “two-body” servo-controls.
A single-body servo-control then has a single body defining one retraction chamber and one extension chamber that are separated by a control piston. The retraction chamber and the extension chamber are then fed from a single hydraulic control valve.
Such a servo-control performs its function well. Nevertheless, for reasons of safety, particularly as from a certain level of force that needs to be developed, the person skilled in the art tends to make use of a two-body servo-control.
A two-body servo-control then has a lower body and an upper body that are assembled in tandem or in parallel.
For example, a servo-control having two bodies in tandem has a slider element with one rod and two pistons, each piston subdividing a corresponding body into a retraction chamber and an extension chamber.
Furthermore, two distinct hydraulic control valves, actuated by a common inlet lever connected to the pilot's flight controls are used for feeding the retraction and extension chambers respectively of the lower and upper bodies.
There also exist servo-controls that are provided with three bodies or even more.
During flight at high speed, extreme maneuvers of the aircraft may give rise to high levels of mechanical stress in the structure of the aircraft. Beyond given load factors, there is the risk of damaging the structure.
In order to warn the pilot that the aircraft is approaching a maneuvering limit, it is possible to provide a device for detecting a limit force on a servo-control. When the force exerted on the servo-control reaches a limit threshold, i.e. a traction force limit or a compression force limit, the limit force detector device triggers a warning, e.g. a visible warning, in order to inform the pilot.
Conventionally, the limit force detector device comprises a detector element provided with a rod fitted with a detection piston that slides in a detection space, the detection space having two detection chambers that are separated by the detection piston, these chambers being independent from the retraction and extension chambers of the outer body. The first detection chamber is fed with fluid by the hydraulic circuit of the aircraft, and the second detection chamber opens out to the outside of the servo-control.
Furthermore, the rod of the detector element projects from the body of the servo-control so as to be connected to the downstream portion of the linkage. This projecting portion of the detector element further includes a lever suitable for co-operating with a pushbutton switch.
Below the limit threshold, the pressure that exists in the detection chamber keeps the detection piston in high abutment so as to ensure that its lever is remote from the switch. In contrast, when the threshold is reached, the pressure that exists in the detection chamber can no longer keep the detection piston in high abutment. The detection piston then reaches a low abutment, with the lever then actuating the switch.
In order to damp the travel of the detection piston in the event of the external force varying at high frequency or in the event of the servo-control being operated quickly, the limit force detection device may include headloss-generator means, e.g. of the diaphragm type, upstream from the first detection chamber.
In order to avoid fluid passing from the detection chamber to the outside of the servo-control, the detection piston carries a gasket. Since this gasket is stressed dynamically, leaks to the outside of the servo-control may nevertheless appear and require maintenance action.
Furthermore, the limit force detector device is subjected to the forces to which the servo-control is subjected by being connected to the control linkage. Under such circumstances, it needs to be dimensioned so as to be capable of withstanding said forces. This gives rise to financial costs and weight that are not negligible.
Finally, the sliding of the detection piston has the practical consequence of giving rise to slack in the control linkage in the event of a drop of pressure in the hydraulic circuit feeding fluid to the limit force detector device.